EP1614622B1 - Aircraft cockpit floor - Google Patents

Abstract

The floor (4) comprises a series of longerons (14) running lengthwise in the aircraft, and cross-members (16) running transversely that make up its primary structure (12), attached to the fuselage by fixings (64). The fixings include rods (68) that take up lengthwise forces, connected by their rear ends to the floor structure by brackets fixed to both longerons and cross-members, and by their forward ends to lengthwise members of the fuselage frame.

Description

TECHNICAL AREA

The present invention relates generally to an aircraft comprising a cockpit floor incorporating in particular a plurality of longitudinal members assembled to a plurality of cross members.

STATE OF THE PRIOR ART

In known manner, the cockpit floor of an aircraft has a shape adapted to the shrinkage of the fuselage encountered in this part of the aircraft, in that its width decreases in going forward.

In addition, this type of floor can extend rearwardly into a cabin portion of the aircraft, and is more generally the floor of the entire nose portion of the aircraft.

Such floors are then designed to meet several specific needs, such as for example the need to present openings for the integration of the control pedals and the central console of the cockpit, allow the march of the occupants of the aircraft, allow the installation of various equipment such as electrical furniture or seats, ensure the recovery of the mechanical forces encountered in case of aircraft crash, or electromagnetically isolate the lower portion of the upper part of this aircraft.

The technological background of an aircraft floor layout is illustrated by the document US Patent 4,479,621 .

From the prior art, there are known cockpit floors incorporating longitudinal members and crosspieces, for example made of aluminum or in one of its alloys, so as to offer good mechanical rigidity. In addition, the overall rigidity of such a floor is enhanced by the presence of boxes obtained by the addition of upper and / or lower metal sheets on a part of the assembly consisting of the longitudinal members and crosspieces.

It is noted that the boxes located at the lateral ends of the floor also serve as means for fixing this floor on the fuselage frames of the cockpit, and therefore allow, thanks to their positioning between the frames, to perform a recovery function efforts acting in the longitudinal direction and transiting between a main structure of the floor and the fuselage frames of the cockpit. As an indication, these forces acting in the longitudinal direction occur mainly in the event of aircraft crash, leading the floor to lean strongly forward against the frames of the fuselage. In addition, this type of effort can also be encountered during the pressurization phases of the aircraft, causing inflation of the cockpit. Nevertheless, the intensity of these efforts is then much lower than that observed during a crash of the aircraft.

On the other hand, the non-boxed parts of the assembly are covered by a honeycomb sandwich type top coating, allowing in particular to the occupants of the aircraft to be able to walk on the floor.

In this type of embodiment of the prior art, it has been found a major drawback from the use of the caissons to ensure the recovery of forces exerted along the longitudinal direction of the aircraft.

Indeed, the assembly of the boxes on the fuselage frames is long and tedious to implement, since it may be necessary to proceed to a successive disassembly and assembly of the lower and upper skins of these boxes.

STATEMENT OF THE INVENTION

The invention therefore aims to propose an aircraft comprising a cockpit floor, this floor at least partially overcoming the disadvantage mentioned above relating to the embodiments of the prior art.

To do this, an aircraft cockpit floor is proposed, this floor comprising a plurality of longitudinal members extending in a longitudinal direction of the aircraft as well as a plurality of cross members assembled to the longitudinal members and extending in a transverse direction of the aircraft. the aircraft, the longitudinal members and cross members constituting a primary structure of the floor, the latter also comprising fixing means enabling it to be assembled on the fuselage of the aircraft. according to According to the invention, the fixing means comprise a plurality of rods for taking up forces acting in the longitudinal direction, these rods each having a rear end connected to the main structure and a front end intended to be connected to the main structure. fuselage, so that each rod is oriented so as to deviate from a center of the floor while going forward.

As an indication, the forward end of the connecting rods is therefore intended to be connected to the fuselage, and preferably to a longitudinal fuselage sail fixedly mounted on the fuselage frames.

Advantageously, the connecting rods of the floor according to the invention ensuring the recovery of the forces exerted in the longitudinal direction can be mounted simply and quickly on the fuselage of the aircraft, proceeding only an assembly of their front end on the same fuselage , for example in an articulated manner. As such, the fact of mounting the front end on a longitudinal fuselage web, fixedly assembled on fuselage frames and preferably on an outer skin of the same fuselage, makes it possible to diffuse the forces in a very homogeneous manner.

It is obvious that the tasks required to assemble such rods are much less restrictive than those for dismantling and assembling successive lower and upper skins of caissons, as was the case previously.

Moreover, the connecting rod solution is entirely adapted to the narrowing of the fuselage width encountered in the cockpit part of the aircraft.

Of course, it should be understood that the rods thus disposed are intended to work in compression in the event of aircraft crash, and to be stressed during the pressurization phases of the aircraft causing inflation of the compartment concerned. However, it is specified that the main role of these rods is to allow the recovery of efforts in case of aircraft crash, the function of recovery of forces exerted in the longitudinal direction during the pressurization phases being so that secondary.

Therefore, such a cockpit floor has the advantage of integrating rods of recovery of longitudinal forces whose length is significantly reduced compared to that of rods which would be oriented to deviate from a center of the floor going towards the rear, because of the narrowing towards the front of the fuselage observed at this cockpit part of the aircraft. The reduction in the length of the connecting rods thus makes it possible to severely limit the buckling problems during the compression phases of these rods, encountered in the event of an aircraft crash.

Preferably, the rear end of each of the connecting rods is connected to one end of one of the longitudinal members and / or to one end of one of the cross members, via a fitting. In such a case where the rear end of each connecting rod is therefore connected to the periphery of the main structure, it can be provided that each fitting associated with a connecting rod is on the one hand fixedly mounted on one end of one of the longitudinal members and / or on one end of one of the crosspieces, and on the other hand hingedly mounted on the rear end of the associated rod. As such, the solution for connecting the rear end of each of the connecting rods at one end of one of the longitudinal members is preferred to that relating to the connection of this rear end to one end of one of the crosspieces, since these are the spars which are dimensioned to correctly recover the forces exerted in the longitudinal direction.

Thus, also providing that the articulation between the fitting and the rear end of the associated rod is oriented in a vertical direction of the aircraft, and that the front end of this rod is mounted on the fuselage via of a hinge also oriented in the vertical direction, it is then clear that each connecting rod is only able to take up the forces exerted in the longitudinal direction of the aircraft and passing between the longitudinal fuselage sail concerned and the main structure of the floor.

Preferably, each fitting is also fixedly mounted on a skin with which it is in contact, this skin covering the main structure of the floor being located between the main structure and the fitting.

In other words, the fitting is secured together with the main structure and the upper skin assembled above it, so that the forces passing through the connecting rod can be transmitted simultaneously to these two elements, then spread through these. In addition, it is clear that connecting the connecting rod to the upper skin allows a better dissemination of the transmitted forces.

As mentioned above, it can be provided that the front end of each of the connecting rods is intended to be mounted in an articulated manner on the fuselage of the aircraft, and preferably on a longitudinal fuselage web mounted on the fuselage frames.

Preferably, each connecting rod is located in a plane defined jointly by the longitudinal and transverse directions, and is inclined with respect to this longitudinal direction by an angle between 0 ° and 30 °. In addition, as mentioned above and because of the narrowing of the width of the fuselage encountered at this level, each rod is then oriented so as to deviate from a center of the floor while going forward, the importance of this spacing being naturally related to the value chosen for the angle mentioned above.

It can be provided that the longitudinal members and the cross members, together forming the primary floor structure, are all made of composite material. This advantageously makes it possible to obtain a significant reduction in the overall mass of this floor. AT As an indication, the reduction in mass observed can go beyond 20% compared to conventional solutions of the prior art using metallic materials.

Moreover, sleepers and spars composite material are advantageously no longer subject to corrosion risks previously encountered.

Finally, it is noted that the type of material used in the floor according to the invention is compatible with all the specific needs mentioned above, especially with regard to the recovery of mechanical forces in case of aircraft crash. .

Preferably, the longitudinal members and the cross members are made of a composite material based on resin-impregnated carbon fibers. This resin is preferably a thermoplastic resin, such as PEEK, PEKK, PPS, etc.

Although PEEK resin is preferred because of the high mechanical performance it is capable of providing, other types of thermoplastic resin could be used, such as the so-called PPS resin mentioned above and obtained by polymerization of the sulfide of phenylene. In addition, thermosetting resins could also be used.

Other advantages and features of the invention will become apparent in the detailed non-limiting description below.

BRIEF DESCRIPTION OF THE DRAWINGS

• la figure 1 représente une vue partiellement éclatée en perspective d'une partie nez d'un aéronef, ce dernier comprenant un plancher de cockpit selon un mode de réalisation préféré de la présente invention ;
• la figure 2 représente une vue en perspective de la structure primaire du plancher de cockpit représenté sur la figure 1 ;
• la figure 3 représente une vue partielle et agrandie en perspective de la figure 2, montrant plus particulièrement l'assemblage entre les sections de longeron et les traverses ;
• la figure 4 représente une vue partielle en perspective du plancher de cockpit montré sur la figure 1, ledit plancher étant représenté sans sa peau ;
• la figure 5 représente une vue de dessus du plancher de cockpit montré sur la figure 1, correspondant au plancher représenté sur la figure 4 auquel il a été assemblé une peau supérieure ainsi que des moyens de fixation permettant son assemblage sur les cadres de fuselage du cockpit ;
• la figure 6 représente une vue en perspective montrant plus spécifiquement les moyens de fixation montrés sur la figure 5 ; et
• la figure 7 représente une vue partielle et agrandie de celle montrée sur la figure 6, à laquelle il a été rajouté une partie du fuselage de l'aéronef.

This description will be made with reference to the appended drawings among which;

• Figure 1 shows a partially exploded perspective view of a nose portion of an aircraft, the latter comprising a cockpit floor according to a preferred embodiment of the present invention;
• Figure 2 shows a perspective view of the primary structure of the cockpit floor shown in Figure 1;
• FIG. 3 represents a partial and enlarged perspective view of FIG. 2, showing more particularly the connection between the spar sections and the crosspieces;
• Figure 4 shows a partial perspective view of the cockpit floor shown in Figure 1, said floor being shown without its skin;
• FIG. 5 represents a top view of the cockpit floor shown in FIG. 1, corresponding to the floor shown in FIG. 4 to which a top skin has been assembled, and fastening means enabling it to be assembled on the fuselage frames of the cockpit. ;
• Figure 6 shows a perspective view showing more specifically the fastening means shown in Figure 5; and
• Figure 7 shows a partial and enlarged view of that shown in Figure 6, to which it has been added a portion of the fuselage of the aircraft.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

With reference to FIG. 1, it is partially represented the front of an aircraft 1, and more precisely the nose part 2 of this aircraft, comprising a cockpit floor 4 according to a preferred embodiment of the present invention.

Throughout the following description, by convention, X is called the longitudinal direction of the aircraft 1, Y is the direction transversely oriented relative to the aircraft, and Z is the vertical direction, these three directions being orthogonal to each other.

On the other hand, the terms "front" and "rear" are to be considered in relation to a direction of advancement of the aircraft encountered following the thrust exerted by the engines of the aircraft, this direction being represented schematically by the arrow 6.

As can be seen in Figure 1, the cockpit floor 4 extends in an XY plane over substantially the entire length of the nose portion 2 of the aircraft, and is mounted on a fuselage 7 thereof. As will be explained in detail later, the cockpit floor 4 is mounted on the one hand on longitudinal fuselage sails 7b, themselves fixedly assembled on fuselage frames 7a and on an outer fuselage skin 7c, and on the other hand fixedly mounted on these same fuselage frames 7a. As an indication, the frames 7a are spaced from each other in the X direction of the aircraft, and distributed on either side of the floor 4 in the Y direction.

In addition, the floor 4 has a shape narrowing in the Y direction while going forward, because of the narrowing of the fuselage 7 in the same direction.

On the other hand, the nose portion 2 may comprise a front space of control station 8 and a cab rear space 10, these two spaces 8, 10 being usually separated by a partition (not shown). More generally, the nose portion of an aircraft and the cockpit floor extend over approximately 10% of the total length of this aircraft in the X direction, that is to say a few meters, and for example on three to five meters. As an illustrative example, when the aircraft is primarily intended to carry cargo and / or military equipment, its nose portion is delimited to the rear by an area for storing the elements that have just been mentioned.

As shown, the cockpit floor 4 can optionally be designed in two separate parts to be mechanically assembled, the separation between a front portion 4a and a rear portion 4b of floor being located for example at the partition separating the front space of the control station 8 and the rear cabin space 10. Nevertheless, to facilitate the understanding of the invention, it will be considered in following the description that the cockpit floor 4 forms a single element extending substantially from one end to the other of the nose portion 2 of the aircraft.

Referring now to Figure 2, there is shown a primary or primary structure 12 of the floor 4 shown in Figure 1, this primary structure 12 being formed by an assembly between a plurality of longitudinal members 14 extending in the direction X, and a plurality of cross members 16 extending in the Y direction of the aircraft. It is specified that this primary structure 12 provides a significant part of the overall stiffness of the cockpit floor 4.

The longitudinal members 14, for example six in number, are each made of a composite material, preferably a thermoplastic composite material obtained by means of folds of carbon fibers impregnated with PEEK, PEKK or PPS resin.

Each spar 14 then preferably has a C-shaped cross section similar to a U-shaped inclined at 90 °, which is particularly easy to obtain with the aid of a stamping press, the latter also making it easy to obtain a C whose upper and lower branches and the base of C have substantially the same thickness, for example between 2 and 5 mm.

In a similar manner, the sleepers 16, for example seven in number, are each made of a composite material, preferably a thermoplastic composite material obtained using folds of carbon fibers impregnated with PEEK, PEKK or PPS.

Each crosspiece 16 then also preferably has a C-shaped cross-section similar to a U-shape inclined at 90 °, and whose upper and lower branches and the base of C have substantially the same thickness, for example between 2 and 5 mm.

Preferably, each crossmember 16 is made in one piece and extends in the Y direction over the entire width of the primary structure 12. On the other hand, each spar 14 is in fact composed of several sections of spar 14a and s'. extends in the X direction over the entire length of the primary structure 12.

More specifically, each spar section 14a of a spar 14 given is positioned between two crosspieces 16 directly consecutive in the direction X, and has two ends rigidly joined respectively to these two same crosspieces 16 directly consecutive.

In this regard, it is noted that the advantage of such a configuration lies in the fact that the upper legs or flanges of the spar sections 14a and cross members 16 are arranged in the same plane XY, these upper branches of C allowing therefore jointly forming a planar top surface of the primary structure 12.

FIG. 3 shows as an indication that the spar sections 14a are assembled to the crosspieces 16 by means of junction elements 20, which are also each made of a composite material, preferably a thermoplastic composite material obtained from the using carbon fiber folds impregnated with PEEK, PEKK or PPS resin.

Overall, each joining element 20 consists of three flat faces which together form a trunk corner. In other words, an element 20 comprises a first plane face 32 which is oriented along an XZ plane, a second plane face 34 which is oriented along a YZ plane, and a third plane face 36 which is oriented in a plane XY, each of these three faces having two junction edges (not referenced) respectively providing the junction with the other two faces. In addition, the three faces 32, 34, 36 preferably have the same thickness, and all meet in a substantially rounded area 37.

Referring now to Figure 4, we see a portion of the cockpit floor 4, the floor 4 comprising the primary structure 12 on which have been assembled the peripheral spars 42, the latter being identical or similar to the spar sections 14a of the primary structure 12. As can clearly be seen in FIG. 4, the peripheral longitudinal members 42 make it possible to connect together and two by two the ends of the cross members 16 of the primary structure 12.

As an illustrative example, it is noted that the floor 4 is also equipped with a small spar 44 located at the rear of the primary structure 12, and defining with a rear cross member 16 a recess 46 in the structure 12, this recess 46 being adapted to accommodate a stair (not shown) of which a top step would be located near the small spar 44.

On the other hand, front secondary beams 48, 49, preferably four in number and made of a thermoplastic composite material obtained using PEEK, PEKK or PPS resin and carbon fiber plies, are secured to the crosses 16 the most front of the primary structure 12.

The two most centrally located secondary beams 48 jointly delimit a space 50 for receiving a central console of the cockpit (not shown), and may each be located in the extension of a spar 14 of the structure 12. In addition, they may be connected to each other at the front via a small cross 51 which can also support the center console.

On the other hand, the two lateral side members 49 each delimit with one of the two secondary side members 48 a space 52 for receiving control pedals (not shown), so that the two spaces 52 obtained are located on both sides. other space 50 in the transverse direction Y of the aircraft.

The cockpit floor 4 furthermore has stiffening elements 54, which preferably extend in the Y direction, between the cross members 16 of the primary structure 12. As an indication, the stiffening elements 54 are made of a composite material, preferably in a thermoplastic composite material obtained using PEEK, PEKK or PPS resin and folds of carbon fibers, and are for example provided in a number between two and five, between two crosspieces 16 directly consecutive in the X direction .

The upper portions of the stiffening elements 54 together define an upper surface which is merged with the upper surface of the primary structure 12, on which a skin 62 such as that shown in FIG. 5 will rest.

This skin 62 is rigidly joined to the longitudinal members 14, the crosspieces 16, as well as to the stiffening elements 54. In this respect, it is specified that these elements 54 are preferably assembled on a lower surface of the skin 62, for example by riveting, before the lower surface of this skin 62 is assembled on the upper branches of the longitudinal members 14 and sleepers 16.

Here again, the skin 62 is preferably made of a composite material with a substantially constant thickness, and preferably in a thermoplastic composite material obtained using PEEK, PEKK or PPS resin and folds of carbon fibers.

In FIG. 5, the primary structure 12 not being covered by a lower skin, it is therefore to be considered that the upper skin 62 forms only half-boxes with the longitudinal members 14 and the crosspieces 16.

Still with reference to this FIG. 5, it can be seen that the floor 4 is equipped with fastening means 64 enabling it to be assembled on the fuselage 7, and more particularly on the longitudinal sails 7b, as well as on the fuselage frames 7a.

Overall, the fastening means 64 are composed of a plurality of articulations 66 allowing a rotation in the X direction, as well as a plurality of connecting rods 68 ensuring the recovery of forces exerted along the same direction X and transiting between the rigid structure 12 and the fuselage 7 (not shown in this figure 5).

Regarding the hinges 66 which will now be briefly described, each of them is mounted at one end of one of the cross members 16, so as to be attached to a fuselage frame 7a located nearby. More specifically, each crosspiece 16 of the floor 4 carries two joints 66 respectively arranged at both ends.

Overall, the presence of such a joint 66 between a cross member 16 and an associated fuselage frame 7a provides a real degree of freedom between these two elements, which has the effect of completely eliminating the moment around of the longitudinal direction X encountered in the embodiments of the prior art, in which such a moment was actually introduced through the fuselage frames 7a and the cross members 16 of the floor 4, mainly during the pressurization phases of the aircraft.

Thus, these joints 66 are therefore able to ensure the recovery of forces exerted along the directions Y and Z and passing between the rigid structure 12 and the fuselage frames 7a of the cockpit.

As for the recovery of forces exerted in the direction X and transiting between the rigid structure 12 and the fuselage 7 of the cockpit, it is provided by the connecting rods 68 which will be presented below.

With reference in conjunction with FIGS. 5 and 6, it can be seen that the fastening means 64 comprise six connecting rods 68, distributed symmetrically with respect to an XZ plane passing through the floor 4 at its center, and thus passing through a longitudinal main axis ( not shown) of the aircraft.

Of course, the number of return rods 68 can be adapted according to the needs encountered, without departing from the scope of the invention.

In this preferred embodiment where each of the two sides of the floor 4 is provided with three connecting rods 68, it is noted that each of these rods 68 is associated with a fitting 70 which is fixedly mounted on the upper skin 62, as well as on a end of one of the longitudinal members 14 and / or on one end of one of the crosspieces 16, the mounting solution on one end of one of the longerons 14 being preferred.

More specifically, each fitting 70 is arranged on the upper skin 62 to which it is secured by means of conventional assembly means of the rivet type, these assembly means thus traversing successively, from top to bottom, the fitting 70, the skin 62, and at least a portion of an end of one of the longitudinal members 14 and / or a portion of one end of one of the crosspieces 16.

Each bracket 70, which is thus fixed to a peripheral portion of the rigid structure 12, extends so as to partially project from the same structure 12 in the X direction and / or in the Y direction, so that this part of the projecting fitting 70 can be easily assembled at a rear end of the associated connecting rod 68.

The assembly is then performed using a hinge 72 arranged between the fitting 70 and the rear end of the associated rod 68, this hinge 72 being oriented in the Z direction of the aircraft, and serving in particular to catch up with construction games.

Similarly, the front end of each connecting rod 68 is intended to be assembled to the longitudinal sail 7b fixedly attached to the fuselage frames 7a and the outer fuselage skin 7c (not shown in FIG. 6). via a hinge 74 oriented in the direction Z, so that these rods 68 are only able to resume the efforts acting in the X direction and transiting between the longitudinal sails 7b and the main structure 12.

Referring more specifically to Figure 7, it will now be described in detail the arrangement between the various elements of the fuselage 7 and the rods 68 of the floor 4.

In this figure 7, we can only see the front left side of the cockpit, but it is understood that the front right side is identical. Therefore, the latter will not be further described.

It can first be seen, as is also shown in Figure 1, that the fuselage 7 comprises the outer skin 7c fixedly mounted on the fuselage frames 7a, which each extend generally in a plane YZ. In addition, the fuselage 7 comprises the longitudinal web 7b which is generally located in an XY plane, preferably substantially identical to that in which is the upper skin 62 covering the main structure 12 of the floor 4.

The longitudinal web 7b, represented truncated in FIG. 7, can be made in several sections, for example two, and is fixedly mounted on the frames 7a as well as on the outer skin 7c, preferably using angles ( not represented). In such a case, it can then be provided that on each side of the cockpit, the longitudinal sail 7b extends substantially parallel to the curvature of the outer skin 7c of the fuselage 7 considered in a plane XY, over a length of about 2.50 m, using two sections of about 1.25 m each.

To obtain a contact and a fixing between an outer edge 7b1 of the longitudinal web 7b and the inner face of the outer skin 7c, the web 7b has recesses 69 allowing the passage of the fuselage frames 7a. Thus, these recesses 69 each take the form of an open notch on the side of the outer edge 7b1 and extending generally in the direction Y, and whose shape is substantially identical to a section of the frame 7a intended to penetrate.

As can be seen in this figure 7, the front end of each connecting rod 68 is connected using a fitting 71 at an inner edge of the longitudinal web 7b. Furthermore, it is stated that in order to ensure a good diffusion of forces, each connecting rod 68 is placed in such a way that its imaginary extension, that is to say its virtual line of sight, intercepts a junction zone between the skin. 7c and a fuselage frame 7a located at the front of the connecting rod 68 concerned.

Referring again to FIG. 6, the positioning of the three fittings 70 respectively associated with the three connecting rods 68 of one of the two sides of the floor 4 will be described in greater detail.

Firstly with regard to the most forward take-up rod 68, the associated fitting 70 is fixed above the skin 62 at a location allowing it to also be secured to one end. before the second outermost beam 14, and the cross 16 the furthest forward.

The intermediate rod 68 is associated with a bracket 70 fixed above the skin 62 at a location that allows it to be also secured to one end of the second cross 16 the furthest forward, and possibly to the one and / or to the other of the two peripheral spars 42 connecting at the end of the crossing that has just been cited. As an indication, this end of crosspiece 16 also carries a hinge 66 oriented in the direction X and belonging to the fixing means 64 of the floor 4, this hinge 66 projecting from the end of the cross member associated in the Y direction, as each of other joints 66.

Finally, the rearmost connecting rod 68 is fixed above the skin 62 at a location that allows it to be secured to a front end of the outermost beam 14 as well as to a substantially central cross member 16, and possibly to the peripheral spar 42 located at this location.

Naturally, as mentioned above, the assembly of the various fittings 70 with the elements just mentioned is preferably carried out using a plurality of rivets oriented in the Z direction, which pass through each associated fitting 70 and at least one of the other elements concerned.

Referring again to FIG. 5, it can be seen that each of the connecting rods 68 is oriented so as to deviate from the center of the floor 4 while going forward, unlike the peripheral beams 42 which deviate from the center of the floor 4 while going backwards.

More specifically, each of the six connecting rods 68, which are preferably all arranged in the same plane XY situated slightly above that of the upper surface of the skin 62, is inclined at an angle α with respect to the direction X which is between 0 ° and 30 °. Preferably, the value of the angle α is chosen to be the closest to 0 °, and may be different for each of the connecting rods 68. Moreover, these rods 68 are each attached to a longitudinal fuselage web 7b, as this has been described with reference to Figure 7.

In addition, it is desirable not to choose rods of too great lengths to avoid problems buckling in compression, or too short lengths that would make difficult mounting adjustment.

Of course, various modifications may be made by those skilled in the cockpit floor 4 which has just been described, only by way of non-limiting example without departing from the scope of the invention defined by the text of the claims.

Claims (11)

1. Aircraft comprising a cockpit floor (4), said floor incorporating a plurality of spars (14) extending along a longitudinal direction (X) of the aircraft and a plurality of cross-beams (16) assembled with said spars (14) and running along a transverse direction (Y) of the aircraft, the spars (14) and the cross-beams (16) forming a primary structure (12) of said floor (14), the floor also comprising attachment means (64) used to assemble it to the aircraft fuselage (7), characterized in that said attachment means (64) comprise a plurality of rods (68) resisting forces applied along the longitudinal direction (X), each of said rods (68) having an aft end connected to the main structure (12) and a forward end that will be connected to said fuselage (12), such that each rod (68) is oriented such that its distance from the centre of the floor (4) increases towards the forward end.
2. Aircraft (4) according to claim 1, characterized in that the forward end will be connected to a longitudinal fuselage wall (7b) fixed directly onto the fuselage frames (7a) of the fuselage (7).
3. Aircraft (4) according to claim 1 or 2, characterized in that the aft end of each of the rods (68) is connected to one end of one of said spars (14) and/or to one end of said cross-beams (16), through a fitting (70).
4. Aircraft (4) according to claim 3, characterized in that each fitting (70) associated with a rod (68) is firstly fixed to one end of one of said spars (14) and/or to one end of one of the cross-beams (16), and secondly articulated on the aft end of the associated rod (68).
5. Aircraft (4) according to claim 4, characterized in that an articulation (72) between the fitting (70) and the aft end of the associated rod (68) is oriented along a vertical direction (Z) of the aircraft.
6. Aircraft (4) according to claim 4 or 5, characterized in that each fitting (70) is also assembled fixed to a skin (62) with which it is in contact, this skin (62) covering the main structure (12) of the floor (4), located between the main structure (12) and the fitting (70).
7. Aircraft (4) according to one of the previous claims, characterized in that the forward end of each of said rods (68) will be assembled articulated on the fuselage (7).
8. Aircraft (4) according to any one of the previous claims, characterized in that each rod (68) is located in a plane defined jointly by the longitudinal (X) and transverse (Y) directions, and in that it is inclined at an angle of between 0° and 30° from said longitudinal direction (X).
9. Aircraft (4) according to any one of the previous claims, characterized in that said spars (14) and said cross-beams (16) are made from a composite material.
10. Aircraft (4) according to claim 9, characterized in that said spars (14) and said cross-beams (16) are made from a thermoplastic composite material.
11. Aircraft (4) according to claim 10, characterized in that the thermoplastic composite material is made using plies of carbon fibres impregnated with PEEK, PEKK or PPS resin.

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